Turbine section of high bypass turbofan

ABSTRACT

A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/793,785, filed Jul. 8, 2015, which is a continuation-in-part of U.S.patent application Ser. No. 14/692,090, filed Apr. 21, 2015, which was acontinuation of U.S. patent application Ser. No. 13/599,175, filed Aug.30, 2012, which was a continuation of U.S. patent application Ser. No.13/475,252, now U.S. Pat. No. 8,844,265, issued Sep. 30, 2014, filed May18, 2012, which was a continuation-in-part of U.S. patent applicationSer. No. 11/832,107, filed Aug. 1, 2007, and claimed the benefit of U.S.Patent Provisional Application No. 61/593,190, filed Jan. 31, 2012, andU.S. Provisional Application No. 61/498,516, filed Jun. 17, 2011.

BACKGROUND

The disclosure relates to turbofan engines. More particularly, thedisclosure relates to low pressure turbine sections of turbofan engineswhich power the fans via a speed reduction mechanism.

There has been a trend toward increasing bypass ratio in gas turbineengines. This is discussed further below. There has generally been acorrelation between certain characteristics of bypass and the diameterof the low pressure turbine section sections of turbofan engines.

SUMMARY

One aspect of the disclosure involves a turbofan engine having an enginecase and a gaspath through the engine case. A fan has a circumferentialarray of fan blades. The engine further has a compressor in fluidcommunication with the fan, a combustor in fluid communication with thecompressor, a turbine in fluid communication with the combustor, whereinthe turbine includes a low pressure turbine section having 3 to 6 bladestages. A speed reduction mechanism couples the low pressure turbinesection to the fan. A bypass area ratio is greater than about 6.0. Aratio of the total number of airfoils in the low pressure turbinesection divided by the bypass area ratio is less than about 170, saidlow pressure turbine section airfoil count being the total number ofblade airfoils and vane airfoils of the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, the bypass area ratio may be greater than about 8.0 or maybe between about 8.0 and about 20.0.

In additional or alternative embodiments of any of the foregoingembodiments, a fan case may encircle the fan blades radially outboard ofthe engine case.

In additional or alternative embodiments of any of the foregoingembodiments, the compressor may comprise a low pressure compressorsection and a high pressure compressor section.

In additional or alternative embodiments of any of the foregoingembodiments, the blades of the low pressure compressor section and lowpressure turbine section may share a low shaft.

In additional or alternative embodiments of any of the foregoingembodiments, the high pressure compressor section and a high pressureturbine section of the turbine may share a high shaft.

In additional or alternative embodiments of any of the foregoingembodiments, there are no additional compressor or turbine sections.

In additional or alternative embodiments of any of the foregoingembodiments, the speed reduction mechanism may comprise an epicyclictransmission coupling the low speed shaft to a fan shaft to drive thefan with a speed reduction.

In additional or alternative embodiments of any of the foregoingembodiments, the low pressure turbine section may have an exemplary 2 to6 blade stages or 2 to 3 blade stages.

In additional or alternative embodiments of any of the foregoingembodiments, a hub-to-tip ratio (Ri:Ro) of the low pressure turbinesection may be between about 0.4 and about 0.5 measured at the maximumRo axial location in the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, a ratio of maximum gaspath radius along the low pressureturbine section to maximum radius of the fan may be less than about0.55, or less than about 0.50, or between about 0.35 and about 0.50.

In additional or alternative embodiments of any of the foregoingembodiments, the ratio of low pressure turbine section airfoil count tobypass area ratio may be between about 10 and about 150.

In additional or alternative embodiments of any of the foregoingembodiments, the airfoil count of the low pressure turbine section maybe below about 1600.

In additional or alternative embodiments of any of the foregoingembodiments, the engine may be in combination with a mountingarrangement (e.g., of an engine pylon) wherein an aft mount reacts atleast a thrust load.

The details of one or more embodiments are set forth in the accompanyingdrawings and the description below. Other features, objects, andadvantages will be apparent from the description and drawings, and fromthe claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an axial sectional view of a turbofan engine.

FIG. 2 is an axial sectional view of a low pressure turbine section ofthe engine of FIG. 1.

FIG. 3 is transverse sectional view of transmission of the engine ofFIG. 1.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

Like reference numbers and designations in the various drawings indicatelike elements.

DETAILED DESCRIPTION

FIG. 1 shows a turbofan engine 20 having a main housing (engine case) 22containing a rotor shaft assembly 23. An exemplary engine is ahigh-bypass turbofan. In such an engine, the normal cruise conditionbypass area ratio of air mass flowing outside the case 22 (e.g., thecompressor sections and combustor) to air mass passing through the case22 is typically in excess of about 4.0 and, more narrowly, typicallybetween about 4.0 and about 12.0. Via high 24 and low 25 shaft portionsof the shaft assembly 23, a high pressure turbine section (gasgenerating turbine) 26 and a low pressure turbine section 27respectively drive a high pressure compressor section 28 and a lowpressure compressor section 30. As used herein, the high pressureturbine section experiences higher pressures that the low pressureturbine section. A low pressure turbine section is a section that powersa fan 42. Although a two-spool (plus fan) engine is shown, one of manyalternative variations involves a three-spool (plus fan) engine whereinan intermediate spool comprises an intermediate pressure compressorbetween the low fan and high pressure compressor section and anintermediate pressure turbine between the high pressure turbine sectionand low pressure turbine section.

The engine extends along a longitudinal axis 500 from a fore end to anaft end. Adjacent the fore end, a shroud (fan case) 40 encircles the fan42 and is supported by vanes 44. An aerodynamic nacelle around the fancase is shown and an aerodynamic nacelle 45 around the engine case isshown.

The low shaft portion 25 of the rotor shaft assembly 23 drives the fan42 through a speed reduction mechanism 46. An exemplary speed reductionmechanism is an epicyclic transmission, namely a star or planetary gearsystem. As is discussed further below, an inlet airflow 520 entering thenacelle is divided into a portion 522 passing along a core flowpath 524and a bypass portion 526 passing along a bypass flowpath 528. With theexception of diversions such as cooling air, etc., flow along the coreflowpath sequentially passes through the low pressure compressorsection, high pressure compressor section, a combustor 48, the highpressure turbine section, and the low pressure turbine section beforeexiting from an outlet 530.

FIG. 3 schematically shows details of the transmission 46. A forward endof the low shaft 25 is coupled to a sun gear 52 (or other high speedinput to the speed reduction mechanism). The externally-toothed sun gear52 is encircled by a number of externally-toothed star gears 56 and aninternally-toothed ring gear 54. The exemplary ring gear is coupled tothe fan to rotate with the fan as a unit.

The star gears 56 are positioned between and enmeshed with the sun gearand ring gear. A cage or star carrier assembly 60 carries the star gearsvia associated journals 62. The exemplary star carrier is substantiallyirrotatably mounted relative via fingers 404 to the case 22.

Another transmission/gearbox combination has the star carrier connectedto the fan and the ring is fixed to the fixed structure (case) ispossible and such is commonly referred to as a planetary gearbox.

The speed reduction ratio is determined by the ratio of diameters withinthe gearbox. An exemplary reduction is between about 2:1 and about 13:1.

The exemplary fan (FIG. 1) comprises a circumferential array of blades70. Each blade comprises an airfoil 72 having a leading edge 74 and atrailing edge 76 and extending from an inboard end 78 at a platform toan outboard end 80 (i.e., a free tip). The outboard end 80 is in closefacing proximity to a rub strip 82 along an interior surface 84 of thenacelle and fan case.

To mount the engine to the aircraft wing 92, a pylon 94 is mounted tothe fan case and/or to the other engine cases. The exemplary pylon 94may be as disclosed in U.S. patent application Ser. No. 11/832,107(US2009/0056343A1). The pylon comprises a forward mount 100 and anaft/rear mount 102. The forward mount may engage the engine intermediatecase (IMC) and the aft mount may engage the engine thrust case. The aftmount reacts at least a thrust load of the engine.

To reduce aircraft fuel burn with turbofans, it is desirable to producea low pressure turbine with the highest efficiency and lowest weightpossible. Further, there are considerations of small size (especiallyradial size) that benefit the aerodynamic shape of the engine cowlingand allow room for packaging engine subsystems.

FIG. 2 shows the low pressure turbine section 27 as comprising anexemplary three blade stages 200, 202, 204. An exemplary blade stagecount is 2-6, more narrowly, 2-4, or 2-3, 3-5, or 3-4. Interspersedbetween the blade stages are vane stages 206 and 208. Each exemplaryblade stage comprises a disk 210, 212, and 214, respectively. Acircumferential array of blades extends from peripheries of each of thedisks. Each exemplary blade comprises an airfoil 220 extending from aninner diameter (ID) platform 222 to an outer diameter (OD) shroud 224(shown integral with the airfoil.

An alternative may be an unshrouded blade with a rotational gap betweenthe tip of the blade and a stationary blade outer air seal (BOAS). Eachexemplary shroud 224 has outboard sealing ridges which seal withabradable seals (e.g., honeycomb) fixed to the case. The exemplary vanesin stages 206 and 208 include airfoils 230 extending from ID platforms232 to OD shrouds 234. The exemplary OD shrouds 234 are directly mountedto the case. The exemplary platforms 232 carry seals for sealing withinter-disk knife edges protruding outwardly from inter-disk spacerswhich may be separate from the adjacent disks or unitarily formed withone of the adjacent disks.

Each exemplary disk 210, 212, 214 comprises an enlarged central annularprotuberance or “bore” 240, 242, 244 and a thinner radial web 246, 248,250 extending radially outboard from the bore. The bore impartsstructural strength allowing the disk to withstand centrifugal loadingwhich the disk would otherwise be unable to withstand.

A turbofan engine is characterized by its bypass ratio (mass flow ratioof air bypassing the core to air passing through the core) and thegeometric bypass area ratio (ratio of fan duct annulus areaoutside/outboard of the low pressure compressor section inlet (i.e., atlocation 260 in FIG. 1) to low pressure compressor section inlet annulusarea (i.e., at location 262 in FIG. 2). High bypass engines typicallyhave bypass area ratio of at least four. There has been a correlationbetween increased bypass area ratio and increased low pressure turbinesection radius and low pressure turbine section airfoil count. As isdiscussed below, this correlation may be broken by having an engine withrelatively high bypass area ratio and relatively low turbine size.

By employing a speed reduction mechanism (e.g., a transmission) to allowthe low pressure turbine section to turn very fast relative to the fanand by employing low pressure turbine section design features for highspeed, it is possible to create a compact turbine module (e.g., whileproducing the same amount of thrust and increasing bypass area ratio).The exemplary transmission is a epicyclic transmission. Alternativetransmissions include composite belt transmissions, metal chain belttransmissions, fluidic transmissions, and electric means (e.g., amotor/generator set where the turbine turns a generator providingelectricity to an electric motor which drives the fan).

Compactness of the turbine is characterized in several ways. Along thecompressor and turbine sections, the core gaspath extends from aninboard boundary (e.g., at blade hubs or outboard surfaces of platformsof associated blades and vanes) to an outboard boundary (e.g., at bladetips and inboard surfaces of blade outer air seals for unshrouded bladetips and at inboard surfaces of OD shrouds of shrouded blade tips and atinboard surfaces of OD shrouds of the vanes). These boundaries may becharacterized by radii R_(I) and R_(O), respectively, which vary alongthe length of the engine.

For low pressure turbine radial compactness, there may be a relativelyhigh ratio of radial span (R_(O)-R_(I)) to radius (R_(O) or R_(I)).Radial compactness may also be expressed in the hub-to-tip ratio(R_(I):R_(O)). These may be measured at the maximum Ro location in thelow pressure turbine section. The exemplary compact low pressure turbinesection has a hub-to-tip ratio close to about 0.5 (e.g., about 0.4-0.5or about 0.42-0.48, with an exemplary about 0.46).

Another characteristic of low pressure turbine radial compactness isrelative to the fan size. An exemplary fan size measurement is themaximum tip radius R_(Tmax) of the fan blades. An exemplary ratio is themaximum R_(O) along the low pressure turbine section to R_(Tmax) of thefan blades. Exemplary values for this ratio are less than about 0.55(e.g., about 0.35-55), more narrowly, less than about 0.50, or about0.35-0.50.

To achieve compactness the designer may balance multiple physicalphenomena to arrive at a system solution as defined by the low pressureturbine hub-to-tip ratio, the fan maximum tip radius to low pressureturbine maximum R_(O) ratio, the bypass area ratio, and the bypass arearatio to low pressure turbine airfoil count ratio. These concernsinclude, but are not limited to: a) aerodynamics within the low pressureturbine, b) low pressure turbine blade structural design, c) lowpressure turbine disk structural design, and d) the shaft connecting thelow pressure turbine to the low pressure compressor and speed reductiondevice between the low pressure compressor and fan. These physicalphenomena may be balanced in order to achieve desirable performance,weight, and cost characteristics.

The addition of a speed reduction device between the fan and the lowpressure compressor creates a larger design space because the speed ofthe low pressure turbine is decoupled from the fan. This design spaceprovides great design variables and new constraints that limitfeasibility of a design with respect to physical phenomena. For examplethe designer can independently change the speed and flow area of the lowpressure turbine to achieve optimal aerodynamic parameters defined byflow coefficient (axial flow velocity/wheel speed) and work coefficient(wheel speed/square root of work). However, this introduces structuralconstraints with respect blade stresses, disk size, material selection,etc.

In some examples, the designer can choose to make low pressure turbinesection disk bores much thicker relative to prior art turbine bores andthe bores may be at a much smaller radius R_(B). This increases theamount of mass at less than a “self sustaining radius”. Another means isto choose disk materials of greater strength than prior art such as theuse of wrought powdered metal disks to allow for extremely highcentrifugal blade pulls associated with the compactness.

Another variable in achieving compactness is to increase the structuralparameter AN² which is the annulus area of the exit of the low pressureturbine divided by the low pressure turbine rpm squared at its redlineor maximum speed. Relative to prior art turbines, which are greatlyconstrained by fan blade tip mach number, a very wide range of AN²values can be selected and optimized while accommodating suchconstraints as cost or a countering, unfavorable trend in low pressureturbine section shaft dynamics. In selecting the turbine speed (andthereby selecting the transmission speed ratio, one has to be mindfulthat at too high a gear ratio the low pressure turbine section shaft(low shaft) will become dynamically unstable.

The higher the design speed, the higher the gear ratio will be and themore massive the disks will become and the stronger the low pressureturbine section disk and blade material will have to be. All of theseparameters can be varied simultaneously to change the weight of theturbine, its efficiency, its manufacturing cost, the degree ofdifficulty in packaging the low pressure turbine section in the corecowling and its durability. This is distinguished from a prior artdirect drive configuration, where the high bypass area ratio can only beachieved by a large low pressure turbine section radius. Because thatradius is so very large and, although the same variables (airfoilturning, disk size, blade materials, disk shape and materials, etc.) aretheoretically available, as a practical matter economics and engine fuelburn considerations severely limit the designer's choice in theseparameters.

Another characteristic of low pressure turbine section size is airfoilcount (numerical count of all of the blades and vanes in the lowpressure turbine). Airfoil metal angles can be selected such thatairfoil count is low or extremely low relative to a direct driveturbine. In known prior art engines having bypass area ratio above 6.0(e.g. 8.0-20), low pressure turbine sections involve ratios of airfoilcount to bypass area ratio above 190.

With the full range of selection of parameters discussed aboveincluding, disk bore thickness, disk material, hub to tip ratio, andR_(O)/R_(Tmax), the ratio of airfoil count to bypass area ratio may bebelow about 170 to as low as 10. (e.g., below about 150 or an exemplaryabout 10-170, more narrowly about 10-150). Further, in such embodimentsthe airfoil count may be below about 1700, or below about 1600.

FIG. 4 shows an embodiment 600, wherein there is a fan drive turbine 608driving a shaft 606 to in turn drive a fan rotor 602. A gear reduction604 may be positioned between the fan drive turbine 608 and the fanrotor 602. This gear reduction 604 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 610 is driven byan intermediate pressure turbine 612, and a second stage compressorrotor 614 is driven by a turbine rotor 216. A combustion section 618 ispositioned intermediate the compressor rotor 614 and the turbine section616.

FIG. 5 shows yet another embodiment 700 wherein a fan rotor 702 and afirst stage compressor 704 rotate at a common speed. The gear reduction706 (which may be structured as disclosed above) is intermediate thecompressor rotor 704 and a shaft 708 which is driven by a low pressureturbine section.

One or more embodiments have been described. Nevertheless, it will beunderstood that various modifications may be made. For example, whenreengineering from a baseline engine configuration, details of thebaseline may influence details of any particular implementation.Accordingly, other embodiments are within the scope of the followingclaims.

The invention claimed is:
 1. A turbofan engine comprising: a fanincluding a circumferential array of fan blades; a compressor in fluidcommunication with the fan, the compressor including a four-stage secondcompressor section and a nine-stage first compressor section, the secondcompressor section including a second compressor section inlet with asecond compressor section inlet annulus area; a fan duct including a fanduct annulus area outboard of the second compressor section inlet,wherein the ratio of the fan duct annulus area to the second compressorsection inlet annulus area defines a bypass area ratio; a combustor influid communication with the compressor; a shaft assembly having a firstportion and a second portion; a turbine in fluid communication with thecombustor, the turbine having a two-stage first turbine section coupledto the first portion of the shaft assembly to drive the first compressorsection, and a four-stage second turbine section coupled to the secondportion of the shaft assembly to drive the fan, each of the secondturbine section including blades and vanes, and a second turbine airfoilcount defined as the numerical count of all of the blades and vanes inthe second turbine section; and a planetary gearbox coupled to the fanand rotatable by the second turbine section through the second portionof the shaft assembly to allow the second turbine to turn faster thanthe fan, the gearbox having a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the gearbox; wherein thesecond turbine airfoil count is below 1600; wherein a ratio of thesecond turbine airfoil count to the bypass area ratio is less than 150,and wherein the second turbine section further includes a maximum gaspath radius and the fan blades include a maximum radius, and a ratio ofthe maximum gas path radius to the maximum radius of the fan blades isequal to or greater than 0.35, and is less than 0.55.
 2. The turbofanengine as recited in claim 1, further comprising a fan case and vanes,the fan case encircling the fan and supported by the vanes.
 3. Theturbofan engine as recited in claim 2, wherein the fan is a single fan,and each fan blade includes a platform and an outboard end having a freetip.
 4. The turbofan engine as recited in claim 3, wherein the gearboxcarries a plurality of gears associated with journals.
 5. The turbofanengine as recited in claim 3, further comprising: an engine aft mountlocation configured to support an engine mount when the engine ismounted and react at least a thrust load of the engine; and an engineforward mount location.
 6. The turbofan engine as recited in claim 5,wherein the engine forward mount location is axially proximate to thegearbox.
 7. The turbofan engine as recited in claim 6, wherein theengine forward mount location engages with an intermediate case.
 8. Theturbofan engine as recited in claim 5, wherein the engine aft mountlocation engages with an engine thrust case.
 9. The turbofan engine asrecited in claim 8, wherein the engine aft mount location is locatedbetween the second turbine section and the first turbine section. 10.The turbofan engine as recited in claim 9 wherein the engine aft mountlocation is located between the second turbine section and the firstturbine section.
 11. The turbofan engine as recited in claim 1, whereinthe second turbine section includes a plurality of blade stagesinterspersed with a plurality of vane stages, and each stage of thesecond turbine section includes a disk with a circumferential array ofblades, each blade including an airfoil extending from an inner diameterto an outer diameter, wherein the inner diameter is associated with aplatform and the outer diameter is associated with a shroud.
 12. Theturbofan engine as recited in claim 11, wherein in at least one stagethe shroud is integral with the airfoil.
 13. The turbofan engine asrecited in claim 12, wherein the shroud includes outboard sealing ridgesconfigured to seal with abradable seals.
 14. The turbofan engine asrecited in claim 13, wherein the abradable seals include honeycomb. 15.The turbofan engine as recited in claim 14, further comprising a caseassociated with the second turbine section, wherein the abradable sealsare fixed to the case.
 16. The turbofan engine as recited in claim 15,wherein each stage of the second turbine section includes a disk, with acircumferential array of blades, each blade including an airfoilextending from an inner diameter to an outer diameter, wherein the innerdiameter is associated with a platform and the outer diameter isunshrouded.
 17. The turbofan engine as recited in claim 15, furthercomprising a stationary blade outer air seal, and a rotational gapbetween the tip and the stationary blade outer air seal.
 18. Theturbofan engine as recited in claim 17, wherein each of the plurality ofvane stages includes a vane, each vane including an airfoil extendingfrom an inner diameter to an outer diameter, wherein the inner diameteris associated with a platform and the outer diameter is associated witha shroud.
 19. The turbofan engine as recited in claim 18, furthercomprising a case associated with the second turbine section, whereinthe shroud is fixed to the case.
 20. The turbofan engine as recited inclaim 19, wherein each platform carries a seal.
 21. The turbofan engineas recited in claim 11, wherein a hub-to-tip ratio (Ri:Ro) of the secondturbine section is between 0.4 and 0.5 measured at the maximum Ro axiallocation in the second turbine section.
 22. The turbofan engine asrecited in claim 1, wherein the ratio of the maximum gas path radius tothe maximum radius of the fan blades is less than 0.50.
 23. The turbofanengine as recited in claim 22, wherein a hub-to-tip ratio (Ri:Ro) of thesecond turbine section is between 0.42-0.48 measured at the maximum Roaxial location in the second turbine section.
 24. A turbofan enginecomprising: a fan including a circumferential array of fan blades; acompressor in fluid communication with the fan, the compressor includinga second compressor section and a first compressor section, the secondcompressor section including a second compressor section inlet with asecond compressor section inlet annulus area; a fan duct including a fanduct annulus area outboard of the second compressor section inlet,wherein the ratio of the fan duct annulus area to the second compressorsection inlet annulus area defines a bypass area ratio; a combustor influid communication with the compressor; a shaft assembly having a firstportion and a second portion; a turbine in fluid communication with thecombustor, the turbine having a first turbine section coupled to thefirst portion of the shaft assembly to drive the first compressorsection, and a second turbine section coupled to the second portion ofthe shaft assembly to drive the fan, each of the second turbine sectionincluding blades and vanes, and a second turbine airfoil count definedas the numerical count of all of the blades and vanes in the secondturbine section; and a planetary gearbox coupled to the fan androtatable by the second turbine section through the second portion ofthe shaft assembly to allow the second turbine to turn faster than thefan, the gearbox having a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the gearbox; wherein thesecond turbine airfoil count is below 1600; wherein a ratio of thesecond turbine airfoil count to the bypass area ratio is less than 150;wherein the second turbine section further includes a maximum gas pathradius and the fan blades include a maximum radius, and a ratio of themaximum gas path radius to the maximum radius of the fan blades is equalto or greater than 0.35, and is less than 0.55, and wherein a hub-to-tipratio (Ri:Ro) of the second turbine section is between 0.4 and 0.5measured at the maximum Ro axial location in the second turbine section;wherein the fist turbine is a two-stage first turbine and the secondturbine is a four-stage second turbine; wherein the first compressor isa nine-stage first compressor; and wherein the second compressor is afour-stage second compressor.
 25. The turbofan engine as recited inclaim 24, further comprising a fan case and vanes, the fan caseencircling the fan and supported by the vanes, and wherein the gearboxcarries a plurality of gears associated with journals.
 26. The turbofanengine as recited in claim 24, further comprising an engine intermediatecase, including an engine forward mount location proximate to thegearbox and configured to support an engine mount when the engine ismounted, and an engine thrust case including an engine aft mountlocation configured to support an engine mount and react at least athrust load when the engine is mounted, wherein the engine aft mountlocation is located between the second turbine section and the firstturbine section.
 27. The turbofan engine as recited in claim 26, whereinthe ratio of the maximum gas path radius to the maximum radius of thefan blades is less than 0.50, and the hub-to-tip ratio (Ri:Ro) isbetween 0.42-0.48.